System for directing airflow into a combustor

ABSTRACT

A system includes a gas turbine combustor. The gas turbine combustor includes a combustion liner disposed about a combustion region and a sleeve disposed about the combustion liner. The combustion liner and the sleeve define an airflow passage circumferentially about the liner. The gas turbine combustor also includes multiple axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor. The multiple axial injectors are asymmetrically configured to provide an uniform injection of the airflow circumferentially about an axis of the gas turbine combustor.

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates to a combustor and, more specifically, to an air injection system for directing airflow into the combustor.

Gas turbine engines typically include a combustor having a combustor liner defining a combustion chamber. Within the combustion chamber, a mixture of compressed air and fuel is combusted to produce hot combustion gases. The combustion gases may flow through the combustion chamber to one or more turbine stages to generate power for driving a load and/or a compressor. Typically, the combustion process heats the combustor liner due to the hot combustion gases. Unfortunately, existing cooling systems to cool the combustor liner may introduce non-uniformity in airflow about the combustion liner and, thus, pressure non-uniformity throughout the combustor. As appreciated, the non-uniform airflow significantly affects the effectiveness of the cooling systems, combustor performance, and emissions.

BRIEF DESCRIPTION OF THE INVENTION

Certain embodiments commensurate in scope with the originally claimed invention are summarized below. These embodiments are not intended to limit the scope of the claimed invention, but rather these embodiments are intended only to provide a brief summary of possible forms of the invention. Indeed, the invention may encompass a variety of forms that may be similar to or different from the embodiments set forth below.

In accordance with a first embodiment, a system includes a gas turbine combustor. The gas turbine combustor includes a combustion liner disposed about a combustion region and a sleeve disposed about the combustion liner. The combustion liner and the sleeve define an airflow passage circumferentially about the liner. The gas turbine combustor also includes multiple axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor. The multiple axial injectors are asymmetrically configured to provide a uniform injection of the airflow circumferentially about an axis of the gas turbine combustor.

In accordance with a second embodiment, a system includes an axial air injection system. The axial air injection system includes multiple axial injectors configured to direct an airflow into an airflow passage defined by a sleeve disposed about a combustion liner surrounding a combustion region of a gas turbine combustor facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor, wherein the airflow passage is disposed circumferentially about the liner. The multiple axial injectors are asymmetrically configured to provide a uniform injection of the airflow circumferentially about an axis of the gas turbine combustor.

In accordance with a third embodiment, a system includes a turbine engine. The turbine engine includes a compressor and a turbine having an axis of rotation, and a combustor coupled to the compressor and the turbine. The combustor includes a combustion liner disposed about a combustion region and a sleeve disposed about the combustion liner, wherein the combustion liner and the sleeve define an airflow passage circumferentially about the liner. The combustor also includes multiple axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor, wherein the multiple axial injectors are configured to increase an effective area for airflow injection further away from the axis of rotation.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:

FIG. 1 is a block diagram of an embodiment of a turbine system having a combustor with an axial air injection system;

FIG. 2 is a cross-sectional view of an embodiment of the turbine system of FIG. 1 having the combustor with the axial air injection system;

FIG. 3 is a partial cross-sectional view of an embodiment of the combustor of FIG. 2, taken within line 3-3, illustrating the axial air injection system;

FIG. 4 is a cross-sectional view of an embodiment of the combustor of FIG. 2, taken along line 4-4, illustrating the axial air injection system (e.g., injectors with different radial heights);

FIG. 5 is a cross-sectional view of an embodiment of the combustor of FIG. 2, taken along line 4-4, illustrating the axial air injection system (e.g., injectors with varying number of struts);

FIG. 6 is a cross-sectional side view of an embodiment of the combustor of FIG. 2 illustrating the axial air injection system (e.g., scoops);

FIG. 7 is a cross-sectional view of an embodiment of the combustor of FIG. 6, taken along line 7-7; and

FIG. 8 is a cross-sectional side view of an embodiment of the combustor of FIG. 2 illustrating an asymmetric airflow passage.

DETAILED DESCRIPTION OF THE INVENTION

One or more specific embodiments of the present invention will be described below. In an effort to provide a concise description of these embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions must be made to achieve the developers' specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure.

When introducing elements of various embodiments of the present invention, the articles “a,” “an,” “the,” and “said” are intended to mean that there are one or more of the elements. The terms “comprising,” “including,” and “having” are intended to be inclusive and mean that there may be additional elements other than the listed elements.

The present disclosure is directed to systems for improving the distribution of airflow into a combustor. In particular, embodiments of the present disclosure include a system that includes the combustor (e.g., gas turbine combustor) with an axial air injection system that includes an asymmetrical configuration to enable a uniform injection of airflow circumferentially about an axis of the combustor. The combustor includes a combustion liner disposed about a combustion region and a sleeve (e.g., flow sleeve) disposed about the combustion liner. The combustion liner and the sleeve define an airflow passage circumferentially about the liner. The axial air injection system includes multiple axial injectors to direct an airflow into the airflow passage thereby facilitating the momentum exchange between an injection flow from the axial air injection system and a cross-flow from an upstream portion of the combustor. The asymmetric configuration of the axial air injection system increases the effectiveness of the momentum exchange. The system may include a turbine engine having an axis of rotation, and the multiple axial injectors increase an effective area for airflow injection further away from the axis of rotation. In certain embodiments, at least two axial injectors include different radial heights. For example, the multiple axial injectors include a first injector further away from the axis of rotation that includes a greater radial height than a second injector closer to the axis of rotation. In some embodiments, at least two axial injectors include different airflow areas. For example, the multiple axial injectors may include an equal radial height, and a first injector further away from the axis of rotation includes a greater effective area for injection of the airflow into the airflow passage than a second injector closer to the axis of rotation. For example, the first injector may include a lesser number of struts through an airflow area than the second injector. In other embodiments, each injector may include an axial scoop disposed on a surface of the sleeve. In further embodiments, the combustion liner and sleeve may each include axes offset from one another to define an asymmetric airflow passage. By utilizing the axial air injection system in the disclosed embodiments, air may be axially injected in a more uniform manner circumferentially about the combustion liner to minimize the pressure drop experienced during combustion due to optimized momentum exchange between the injection flow and crossflow, while improving the flame holding margin (e.g., reducing the possibility of flame holding). The more uniform airflow also enhances the cooling of the combustion liner and the combustor performance, while reducing emissions.

FIG. 1 is a block diagram of an embodiment of a gas turbine system 10 (e.g., gas turbine engine) that employs an asymmetrical axial air injection system that includes a plurality of axial injectors. The gas turbine system 10 includes a combustor 12 (e.g., gas turbine combustor) that combusts fuel 14 to drive the gas turbine system 10. The gas turbine system 10 may include one or more combustors 12. As described in detail below, the combustor 14 includes a combustor liner disposed about a combustion region and a sleeve (e.g., flow sleeve) disposed about the combustion liner. The combustion liner and the sleeve define an airflow passage circumferentially about the liner. By utilizing the axial air injection system in the disclosed embodiments, air may be axially injected in a more uniform manner circumferentially about the combustion liner to minimize the pressure drop experienced during combustion, while improving the flame holding margin. The more uniform airflow also enhances the cooling of the combustion liner and the combustor performance, while reducing emissions. In particular, the asymmetrically configured plurality of axial injectors directs an airflow into the airflow passage circumferentially about the liner. The plurality of axial injectors increases an effective area for airflow injection further away from an axis of rotation of the turbine engine 10 (e.g., on the outboard or low pressure side of the sleeve). In particular, the plurality of axial injectors provides a uniform injection of airflow circumferentially about an axis of the combustor 12.

Within the combustor 12, the fuel 14 may mix with compressed air 16, shown by arrows, and ignition may occur, producing hot combustion gases 18 that power the gas turbine system 10. According to certain embodiments, the fuel 14 may be a liquid or gaseous fuel, such as natural gas, light or heavy distillate oil, naphtha, crude oil, residual oil, or syngas. The compressed air 16 includes intake air 20 that enters the gas turbine system 10 through an air intake section 22. The intake air 20 is compressed by a compressor 24 to produce the compressed air 16 that enters the combustor 12. In certain embodiments, one or more fuel nozzles may direct the fuel 14 and/or the compressed air 16 into the combustion region of the combustor 12. In addition, the axial air injection system may direct compressed air 16 into the combustion region. Within the combustion zone, the compressed air 16 combusts with the fuel 14 to produce the hot combustion gases 18. From the combustor 12, the hot combustion gases 18 may flow through a turbine 26 that drives the compressor 24 via a shaft 28. For example, the combustion gases 18 may apply motive forces to turbine rotor blades within the turbine 26 to rotate the shaft 28. Shaft 28 also may be connected to a load 30, such as a generator, a propeller, a transmission, or a drive system, among others. After flowing through the turbine 26, the hot combustion gases 18 may exit the gas turbine system 10 through an exhaust section 32.

FIG. 2 is a cross-sectional view of an embodiment of the turbine system 10 (e.g., gas turbine) of FIG. 1 having the combustor 12 with the axial air injection system 42. In certain embodiments, the gas turbine engine 10 may include multiple combustors 12 disposed in an annular arrangement. As described above with respect to FIG. 1, air enters the gas turbine engine 10 through the air intake 22 and is pressurized in the compressor 24. Compressed air and fuel are directed through an end cover 44 and a head end 46 to each of the fuel nozzles 47, which distribute a fuel-air mixture into the combustor 12. The combustor 12 includes a combustion chamber 48 (e.g., combustion region), which is generally defined by a combustion casing 50, a combustion liner 52, and a sleeve 54 (e.g., flow sleeve). In certain embodiments, the sleeve 54 and the combustion liner 52 are coaxial with one another to define a hollow annular space 49 (e.g., airflow passage), which may enable passage of air for cooling and for entry into the head end 46 and the combustion chamber 48. In particular, the combustor 12 includes the combustion liner 52 disposed about the combustion region 48 and the sleeve 54 disposed about the combustion liner 52. The combustion liner 52 and the sleeve define the airflow passage 49 circumferentially 55 about the liner 52.

The axial air injection system 42 axially injects air into the airflow passage 49 in a more uniform manner circumferentially 55 about the combustion liner 52 (e.g., about an axis 57 of the combustor 12) minimizing the pressure drop experienced during combustion, while improving the flame holding margin. The more uniform airflow also enhances the cooling of the combustion liner 52 and the combustor performance, while reducing emissions. In particular, asymmetrically configured axial injectors direct an airflow generally in axial direction 59 into the airflow passage 49 circumferentially 55 about the liner 52. The plurality of axial injectors increases an effective area for airflow injection further away from an axis of rotation 58 of the turbine engine 10 (e.g., on the outboard or low pressure side of the sleeve 54). In particular, the plurality of axial injectors provides a uniform injection of airflow circumferentially 55 about the axis 57 of the combustor 12.

The design of the combustor 12 provides optimal flow of the air-fuel mixture through a transition piece 60 (e.g., converging section) towards the turbine 26. For example, the fuel nozzles 47 may distribute the pressurized air-fuel mixture into the combustion chamber 48, where combustion of the air-fuel mixture occurs. The resultant exhaust gas flows through the transition piece 60 to the turbine 26, as illustrated by arrow 62, causing the blades 64 of the turbine 26 to rotate, along with the shaft 28, about the axis of rotation 58. The rotation of the turbine blades 64 causes a rotation of the shaft 28, thereby causing blades 66 within the compressor 24 to draw in and pressurize the air received by the intake 22.

FIG. 3 is a partial cross-sectional view of an embodiment of the combustor 12 of FIG. 2, taken within line 3-3, illustrating the axial air injection system 42. The combustor 12 includes the combustion liner 52, the sleeve 54 (e.g., flow sleeve), and the axial air injection system 42. As mentioned above, the combustion liner 52 and the sleeve 54 define the airflow passage 49 circumferentially 55 about the liner 52. In certain embodiments, the combustion liner 52 and the sleeve 54 each include axes offset from one another to define asymmetry in the airflow passage 49 (see FIG. 8). The axial air injection system 42 includes a plurality of axial injectors 76 disposed adjacent an end 78 of the sleeve 54 opposite the head end 46 of the combustor 12. The plurality of axial injectors 76 is disposed circumferentially 55 about the end 78 of the sleeve 54 and extends partially into the airflow passage 49 in radial direction 80. The plurality of axial injectors 76 are configured to direct airflow (e.g., compressed air from the compressor 24) into the airflow passage 49 generally in axial direction 59. For example, each injector 76 includes an opening 82 for directing airflow in the axial direction 59.

As illustrated, compressed air from the compressor 24 flows generally in axial direction 59 outside of the sleeve 54 as generally indicated by arrow 84. Also, compressed air flows generally in axial direction 59 into the airflow passage 49 from the transition piece 60 as generally indicated by arrow 85. A lip 86 of the plurality of axial injectors 76 extends beyond the sleeve 54 in radial direction 88, such that the lip 86 directs (e.g., turns) the airflow outside the sleeve 54 from the axial direction 59 to the radial direction 80 as generally indicated by arrow 90. Then, the plurality of axial injectors 76 directs the airflow generally in axial direction 59 through the openings 82 as generally indicated by arrow 92. Injector airflow 94 entrains (e.g., facilitating a momentum exchange) the airflow 85 (e.g., crossflow) from the transition piece 60 (e.g., upstream portion of the combustor 12) into a single airflow 96 generally in axial direction 59 toward the head end 46 of the combustor 12.

As described in greater detail below, the plurality of axial fuel injectors 76 include an asymmetric configuration to provide a uniform injection of airflow circumferentially 55 about the axis 57 of the combustor 12. In particular, the asymmetric configuration of the axial fuel injectors 76 increases the effective area for airflow injection further away from the axis 58 of rotation of the turbine engine 10 (e.g., on the outboard or low pressure side of the sleeve 54). This enables the mass flow and velocity of the air injected into the airflow passage 49 on the outboard side to match the mass flow and velocity of the air injected on the inboard or high pressure side of the sleeve 54 to create the uniform airflow circumferentially 55 about the axis 57 of the combustor 12. In certain embodiments, at least two axial injectors 76 of the plurality of axial injectors 76 include different radial heights. For example, at least one injector 76 further away from the axis of rotation 58 of the turbine engine 10 includes a greater radial height than another injector 76 closer to the axis of rotation 58. In other embodiments, the plurality of axial injectors 76 include an equal radial height, and the at least one injector 76 further away from the axis of rotation 58 includes a greater effective area for injection of the airflow into the airflow passage 49 than another injector 76 closer to the axis of rotation 58. For example, each injector 76 of the plurality of axial injectors 76 includes an airflow area, and at least two axial injectors 76 of the plurality of axial injectors 76 include different airflow areas for injection of the airflow into the airflow passage 49. For example, at least one injector 76 has a lesser number of struts through the airflow area than another injector 76. In some embodiments, the plurality of axial injectors 76 includes axial scoops disposed on the surface of the sleeve 54 (see FIGS. 7 and 8). By utilizing the axial air injection system 42, air may be axially injected in a more uniform manner circumferentially 55 about the combustion liner 52 minimizing the pressure drop experienced during combustion, while improving the flame holding margin. The more uniform airflow also enhances the cooling of the combustion liner 52 and the combustor performance, while reducing emissions.

FIGS. 4-8 illustrate different embodiments of the axial air injection system 42. Each embodiment of the axial air injection system 42 enables air to be axially injected in a more uniform manner circumferentially 55 about the combustion liner 52 minimizing the pressure drop experienced during combustion, while improving the flame holding margin. The more uniform airflow also enhances the cooling of the combustion liner 52 and the combustor performance, while reducing emissions. FIG. 4 is a cross-sectional view of an embodiment of the combustor 12 of FIG. 2, taken along line 4-4, illustrating the axial air injection system 42 having injectors 76 with different radial heights. The number of injectors 76 in the axial air injection system 42 may vary from 1 to 50, 1 to 5, 1 to 10, 1 to 20, 10 to 20, 20 to 30, 30 to 40, or 40 to 50, or any other number. As illustrated, the combustor 12 includes the axial air injection system 42 with an asymmetrical configuration. In particular, the axial air injection system 42 includes the plurality of axial injectors 76 having an asymmetrical configuration to provide a uniform injection of airflow into the airflow passage 49 circumferentially 55 about the axis 57 of the combustor 12. In particular, the injectors 76 direct airflow generally in axial direction 59 into the airflow passage 49. The asymmetric configuration of the axial fuel injectors 76 increases the effective area for airflow injection further away from the axis 58 of rotation of the turbine engine 10 (e.g., on the outboard or low pressure side 106 of the sleeve 54). This enables the mass flow and velocity of the air injected into the airflow passage 49 on the outboard side to match the mass flow and velocity of the air injected on the inboard or high pressure side 108 of the sleeve 54 to create the uniform airflow circumferentially 55 about the axis 57 of the combustor 12.

Each axial injector 76 includes a radial height 110 and an airflow area 111 for the openings 82. In certain embodiments, at least two axial injectors 76 of the plurality of axial injectors 76 include different radial heights 110. For example, at least one injector 76 further away from the axis of rotation 58 (e.g., on outboard side 106) of the turbine engine 10 includes a greater radial height 110 than another injector 76 closer to the axis of rotation 58 (e.g., on inboard side 108). As illustrated, radial heights 110 of axial injectors 112, 114, and 116 differ from the radial heights 110 of axial injectors 118 and 120. Specifically, the radial heights 110 of injectors 112, 114, and 116 are greater than the radial heights 110 of injectors 118 and 120. The radial heights 110 between the injectors 76 may vary (e.g., the height 110 of one injector 76 may be greater than another injector 76) between each other (e.g., between injectors 112, 114, 116 and injectors 118 and 120) from approximately 1 to 200 percent, 1 to 100 percent, 10 to 50 percent, 50 to 100 percent, 100 to 200 percent, 100 to 150 percent, 150 to 200 percent, and all subranges therebetween. For the example, the radial heights 110 between the injectors 76 may vary (e.g., the height 110 of one injector 76 may be greater than another injector 76) by approximately 10, 20, 30, 40, 50, 60, 70, 80, 90, 100, 110, 120, 130, 140, 150, 160, 170, 180, 190, or 200 percent, or any other number.

In certain embodiments, at least two axial injectors 76 of the plurality of axial injectors 76 include different airflow areas 111 (e.g., both geometric and effective areas for airflow). For example, at least one injector 76 further away from the axis of rotation 58 (e.g., on outboard side 106) of the turbine engine 10 includes a greater airflow area 111 than another injector 76 closer to the axis of rotation 58 (e.g., on inboard side 108). As illustrated, the airflow areas 111 of axial injectors 112, 114, and 116 differ from the airflow areas 111 of axial injectors 118 and 120. Specifically, the airflow areas 111 of injectors 112, 114, and 116 are greater than the airflow areas 111 of injectors 118 and 120. The airflow areas 111 between the injectors 76 may vary (e.g., the area 111 of one injector 76 may be greater than another injector 76) between each other (e.g., between injectors 112, 114, 116 and injectors 118 and 120) from approximately 1 to 200 percent, 1 to 100 percent, 10 to 50 percent, 50 to 100 percent, 100 to 200 percent, 100 to 150 percent, 150 to 200 percent, and all subranges therebetween. For the example, the airflow areas 111 between the injectors 76 may vary (e.g., the area 111 of one injector 76 may be greater than another injector 76) by approximately 10, 20, 30, 40, 50, 60, 70, 80, 90, 100, 110, 120, 130, 140, 150, 160, 170, 180, 190, or 200 percent, or any other number.

Alternatively, the axial injectors 76 may have the same geometric airflow areas 11 (e.g., area to outer perimeter), but different effective airflow areas 111 (e.g., area to outer perimeter less any intermediate obstructions). FIG. 5 is a cross-sectional view of an embodiment of the combustor of FIG. 2, taken along line 4-4, illustrating the axial air injection system 42 having injectors 76 with varying number of struts. The number of injectors 76 in the axial air injection system 42 may vary from 1 to 50, 1 to 5, 1 to 10, 1 to 20, 10 to 20, 20 to 30, 30 to 40, or 40 to 50, or any other number. As illustrated, the combustor 12 includes the axial air injection system 42 with an asymmetrical configuration. In particular, the axial air injection system 42 includes the plurality of axial injectors 76 having an asymmetrical configuration to provide a uniform injection of airflow into the airflow passage 49 circumferentially 55 about the axis 57 of the combustor 12. In particular, the injectors 76 direct airflow generally in axial direction 59 into the airflow passage 49. The asymmetric configuration of the axial fuel injectors 76 increases the effective area for airflow injection further away from the axis 58 of rotation of the turbine engine 10 (e.g., on the outboard or low pressure side 106 of the sleeve 54). This enables the mass flow and velocity of the air injected into the airflow passage 49 on the outboard side to match the mass flow and velocity of the air injected on the inboard or high pressure side 108 of the sleeve 54 to create the uniform airflow circumferentially 55 about the axis 57 of the combustor 12.

As illustrated, the plurality of axial injectors 76 includes an equal radial height 110 and an equal geometric airflow area 111 (e.g., area to outer perimeter). However, in certain embodiments, at least two axial injectors 76 of the plurality of axial injectors 76 include different effective airflow areas 111 (e.g., area to outer perimeter less any intermediate obstructions) due to the presence or absence of struts 128 disposed within the openings 82 of the injectors 76. For example, at least one injector 76 further away from the axis of rotation 58 (e.g., on outboard side 106) of the turbine engine 10 includes a greater effective airflow area 111 than another injector 76 closer to the axis of rotation 58 (e.g., on inboard side 108). As illustrated, the effective airflow areas 111 of axial injectors 130 and 132 differ from the effective airflow areas 111 of axial injectors 134, 136, and 138. Specifically, the effective airflow areas 111 of injectors 130 and 132 are greater than the effective airflow areas 111 of injectors 134, 136, and 138. Also, the effective airflow areas 111 of injectors 134 and 138 are greater than the effective airflow area 11 of injector 136. The airflow areas 111 between the injectors 76 may vary (e.g., the area 111 of one injector 76 may be greater than another injector 76) between each other (e.g., between injectors 130 and 132 and injectors 134, 136, and 138) from approximately 1 to 200 percent, 1 to 100 percent, 10 to 50 percent, 50 to 100 percent, 100 to 200 percent, 100 to 150 percent, 150 to 200 percent, and all subranges therebetween. For the example, the effective airflow areas 111 between the injectors 76 may vary (e.g., the area 111 of one injector 76 may be greater than another injector 76) by approximately 10, 20, 30, 40, 50, 60, 70, 80, 90, 100, 110, 120, 130, 140, 150, 160, 170, 180, 190, or 200 percent, or any other number.

As mentioned above, the variance in effective airflow area 76 is due to the presence or absence of struts 128 in the openings 82 of the injectors 76. The number of struts 128 decreases from the inboard side 108 to the outboard side 106 (i.e., away from the axis of rotation 58 of the turbine engine 10). In certain embodiments, at least two injectors 76 of the plurality of axial injectors 76 include a different number of struts 128. For example, at least one injector 76 has a lesser number of struts 128 through the airflow area 111 than another injector 76. Specifically, injectors 130 and 132 (e.g., 0 struts 128) include a lesser number of struts 128 than injectors 134, 136, and 138 (e.g., 2 to 3 struts 128). Also, injectors 134 and 138 (e.g., 2 struts 128) include a lesser number of struts 128 than injector 136 (e.g., 3 struts). The number of struts 128 may vary from 1 to 10, 1 to 5, or 1 to 3, or any other number.

In certain embodiments, the axial injectors 76 may include an axial scoop as illustrated in FIGS. 6 and 7. FIG. 6 is a cross-sectional side view of an embodiment of the combustor 12 of FIG. 2 illustrating the axial air injection system 42 having scoops 148 (e.g., axial scoops). FIG. 7 is a cross-sectional view of the combustor 12 of FIG. 6, taken along line 7-7. The number of injectors 76 (e.g., scoops 148) in the axial air injection system 42 may vary from 1 to 200, 1 to 50, 20 to 40, 50 to 100, 70 to 90, 100 to 150, 120 to 140, 150 to 200, 170 to 190, or any other number. The scoops 148 are disposed on a surface 150 of the sleeve 54. In particular, the scoops 148 are circumferentially 55 disposed about the surface 150 of the sleeve 54. As illustrated, the combustor 12 includes the axial air injection system 42 with an asymmetrical configuration. In particular, the axial air injection system 42 includes the plurality of axial injectors 76 (e.g., scoops 148) having an asymmetrical configuration to provide a uniform injection of airflow circumferentially 55 about the axis 57 of the combustor 12. In particular, the scoops 148 direct airflow generally in axial direction 59 into the airflow passage 49. The scoops 148 each include radial heights 110. In certain embodiments, at least two scoops 148 of the plurality of axial injectors 76 include different radial heights 110. For example, at least one scoop 148 further away from the axis of rotation 58 (e.g., on outboard side 106) of the turbine engine 10 includes a greater radial height 110 than another scoop 148 closer to the axis of rotation 58 (e.g., on inboard side 108). As illustrated in FIG. 6, radial heights 110 of scoops 152 and 154 differ. In particular, the radial height 110 of scoop 152 is greater than the radial height 110 of scoop 154. As depicted in both FIGS. 6 and 7, the radial heights 110 of the scoops 148 may decrease circumferentially 55 about the sleeve 54 from the outboard side 106 to the inboard side 108. The radial heights 110 between the scoops 148 may vary (e.g., the height 110 of one injector 76 may be greater than another injector 76) between each other (e.g., betweens scoops 152 and 154) from approximately 1 to 200 percent, 1 to 100 percent, 10 to 50 percent, 50 to 100 percent, 100 to 200 percent, 100 to 150 percent, 150 to 200 percent, and all subranges therebetween. For the example, the radial heights 110 between the scoops 148 may vary (e.g., the height 110 of one injector 76 may be greater than another injector 76) by approximately 10, 20, 30, 40, 50, 60, 70, 80, 90, 100, 110, 120, 130, 140, 150, 160, 170, 180, 190, or 200 percent, or any other number.

In certain embodiments, at least two scoops 148 of the plurality of axial injectors 76 include different airflow areas 111 (e.g., both geometric and effective areas for airflow) as illustrated in FIG. 7. For example, at least scoop 148 further away from the axis of rotation 58 (e.g., on outboard side 106) of the turbine engine 10 enables includes a greater airflow area 111 than another scoop 148 closer to the axis of rotation 58 (e.g., on inboard side 108). As illustrated, the airflow areas 111 of scoop 152 differs from the airflow areas 111 of scoop 154. Specifically, the airflow areas 111 of scoop 152 is greater than the airflow areas 111 of scoop 154. The airflow areas 111 between the scoops 154 may vary (e.g., the area 111 of one injector 76 may be greater than another injector 76) between each other (e.g., between scoops 152 and 154) from approximately 1 to 200 percent, 1 to 100 percent, 10 to 50 percent, 50 to 100 percent, 100 to 200 percent, 100 to 150 percent, 150 to 200 percent, and all subranges therebetween. For the example, the airflow areas 111 between the scoops 148 may vary (e.g., the area 111 of one injector 76 may be greater than another injector 76) by approximately 10, 20, 30, 40, 50, 60, 70, 80, 90, 100, 110, 120, 130, 140, 150, 160, 170, 180, 190, or 200 percent, or any other number.

The asymmetric configuration (e.g., due to the differences in radial heights 110 and airflow areas 111) of the plurality of axial fuel injectors 76 (e.g., scoops 148) increases the effective area for airflow injection further away from the axis 58 of rotation of the turbine engine 10 (e.g., on the outboard or low pressure side 106 of the sleeve 54). This enables the mass flow and velocity of the air injected into the airflow passage 49 on the outboard side 106 to match the mass flow and velocity of the air injected on the inboard or high pressure side 108 of the sleeve 54 to create the uniform airflow circumferentially 55 about the axis 57 of the combustor 12.

As illustrated, compressed air from the compressor 24 flows generally in axial direction 59 outside of the sleeve 54 as generally indicated by arrow 156. Also, compressed air flows generally in axial direction 59 into the airflow passage 49 from the transition piece 60 as generally indicated by arrow 158. The scoops 148 radially 80 and 88 extend beyond the sleeve 54 to direct the airflow 156 outside the sleeve 54 into the airflow passage 49 through openings 159, while generally maintaining the airflow in the axial direction 59 as generally indicated by arrows 160. Injector airflow 160 entrains (e.g., facilitating a momentum exchange) the airflow 158 (e.g., crossflow) from the transition piece 60 (e.g., upstream portion of the combustor 12) into a single airflow 162 generally in axial direction 59 toward the head end 46 of the combustor 12.

As mentioned above, in certain embodiments the airflow passage 49 itself may be asymmetric. FIG. 8 is a cross-sectional side view of an embodiment of the combustor 12 of FIG. 2 illustrating an asymmetric airflow passage 49. The combustor 12 includes the sleeve 54 (e.g., flow sleeve) and the combustion liner 52. As mentioned above, the combustor 12 includes the sleeve 54 disposed about the combustion liner 52. The combustion liner 52 and the sleeve 54 define the airflow passage 49 circumferentially 55 about the liner 52. The combustion liner 52 (and the combustor 12) include the axis 57 and the sleeve 54 includes the axis 172. The axes 57 and 172 are offset from one another to define asymmetry in the airflow passage 49. In particular, the axes 57 and 172 are offset by a distance 173. Due to the offset distance 173, the airflow passage 49 includes a greater radial height 174 away from the axis of rotation 58 (e.g., on outboard side 106) of the turbine engine 10 than the radial height 174 of the airflow passage 49 closer to the axis of rotation 58 (e.g., on inboard side 108). The radial height 174 of the airflow passage 49 near the outboard side 106 and the radial height 174 of the airflow passage 49 near the inboard side 108 may vary between each other from approximately 1 to 200 percent, 1 to 100 percent, 10 to 50 percent, 50 to 100 percent, 100 to 200 percent, 100 to 150 percent, 150 to 200 percent, and all subranges therebetween.

The asymmetric configuration (e.g., due to the offset of axes 57 and 172) of the airflow passage 49 increases the effective area for airflow injection further away from the axis 58 of rotation of the turbine engine 10 (e.g., on the outboard or low pressure side 106 of the sleeve 54). This enables the mass flow and velocity of the air injected into the airflow passage 49 on the outboard side 106 to match the mass flow and velocity of the airflow injected on the inboard or high pressure side 108 of the sleeve 54 to create the uniform airflow circumferentially 55 about the axis 57 of the combustor 12. The uniform airflow minimizes the pressure drop experienced during combustion, while improving the flame holding margin. The more uniform airflow also enhances the cooling of the combustion liner 52 and the combustor performance, while reducing emissions.

Technical effects of the disclosed embodiments include providing systems to improve the distribution of airflow into the combustor 12. In particular, the combustor includes the axial air injection system 42 that includes an asymmetrical configuration to enable an uniform injection of airflow into the airflow passage 49 between the sleeve 54 and the combustion liner 53 circumferentially about an axis 57 of the combustor 57. The axial air injection system 42 includes multiple axial injectors 76 to direct an airflow axially into the airflow passage 49. The axial injectors 76 may include different radial heights 110 and/or effective airflow areas 111 to increase the effective area for inflow injection further away from the axis 58 of rotation of the turbine engine 10. In certain embodiments, the injectors 76 may include axial scoops 148 disposed on the surface 150 of the sleeve 54. In further embodiments, the combustion liner 52 and sleeve 54 may each include axes 57 and 172 offset from one another to define an asymmetric airflow passage 49. By utilizing the axial air injection system 42, air may be axially injected into the airflow passage 49 in a more uniform manner circumferentially about the combustion liner 52 minimizing the pressure drop experienced during combustion, while improving the flame holding margin. The more uniform airflow also enhances the cooling of the combustion liner 52 and the combustor performance, while reducing emissions.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims. 

1. A system, comprising: a gas turbine combustor, comprising: a combustion liner disposed about a combustion region; a sleeve disposed about the combustion liner, wherein the combustion liner and the sleeve define an airflow passage circumferentially about the liner; and a plurality of axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor, wherein the plurality of axial injectors is asymmetrically configured to provide a uniform injection of the airflow circumferentially about an axis of the gas turbine combustor.
 2. The system of claim 1, wherein the system comprises a turbine engine having an axis of rotation, and the plurality of axial injectors is configured to increase an effective area for airflow injection further away from the axis of rotation.
 3. The system of claim 2, wherein a first injector of the plurality of axial injectors further away from the axis of rotation comprise a greater radial height than a second injector of the plurality axial injectors closer to the axis of rotation.
 4. The system of claim 2, wherein the plurality of axial injectors comprises an equal radial height, and a first injector of the plurality of axial injectors further away from the axis of rotation comprises a greater effective area for injection of the airflow into the airflow passage than a second injector of the plurality of axial injectors closer to the axis of rotation.
 5. The system of claim 4, wherein each injector of the plurality of axial injectors comprises an airflow area, and the first injector has a lesser number of struts through the airflow area than the second injector.
 6. The system of claim 1, wherein at least two axial injectors of the plurality of axial injectors comprise different radial heights.
 7. The system of claim 1, wherein at least two axial injectors of the plurality of axial injectors comprise different airflow areas for injection of the airflow into the airflow passage.
 8. The system of claim 1, wherein each injector of the plurality of axial injectors comprises an axial scoop disposed on a surface of the sleeve.
 9. The system of claim 1, wherein the combustor liner comprises a first axis and the sleeve comprises a second axis, and the first and second axes are offset from one another to define asymmetry in the airflow passage.
 10. The system of claim 1, wherein the system comprises a gas turbine engine having the gas turbine combustor.
 11. A system, comprising: an axial air injection system, comprising: a plurality of axial injectors configured to direct an airflow into an airflow passage defined by a sleeve disposed about a combustion liner surrounding a combustion region of a gas turbine combustor facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor, wherein the airflow passage is disposed circumferentially about the liner, and the plurality of axial injectors is asymmetrically configured to provide a uniform injection of the airflow circumferentially about an axis of the gas turbine combustor.
 12. The system of claim 11, wherein each injector of the plurality of axial injectors comprises an airflow area, and a first injector of the plurality of axial injectors has a lesser number of struts through the airflow area than a second injector of the plurality of axial injectors.
 13. The system of claim 11, wherein at least two axial injectors of the plurality of axial injectors comprise different radial heights.
 14. The system of claim 11, wherein at least two axial injectors of the plurality of axial injectors comprise different airflow areas for injection of the airflow into the airflow passage.
 15. The system of claim 11, wherein the system comprises a turbine engine having an axis of rotation, and the plurality of axial injectors is configured to increase an effective area for airflow injection further away from the axis of rotation.
 16. The system of claim 15, wherein a first injector of the plurality of axial injectors further away from the axis of rotation comprises a greater radial height than a second injector of the plurality of axial injectors closer to the axis of rotation.
 17. A system, comprising; a turbine engine, comprising: a compressor and a turbine having an axis of rotation; and a combustor coupled to the compressor and the turbine, wherein the combustor comprises: a combustion liner disposed about a combustion region; a sleeve disposed about the combustion liner, wherein the combustion liner and the sleeve define an airflow passage circumferentially about the liner; and a plurality of axial injectors configured to direct an airflow into the airflow passage in an axial direction facilitating a momentum exchange between an injection flow and a crossflow from an upstream portion of the combustor, wherein the plurality of axial injectors is configured to increases an effective area for airflow injection further away from the axis of rotation.
 18. The system of claim 17, wherein the plurality of axial injectors is asymmetrically configured to provide a uniform injection of the airflow circumferentially about an axis of the combustor.
 19. The system of claim 17, wherein a first injector of the plurality of the plurality of axial injectors further away from the axis of rotation comprises a greater radial height than a second injector of the plurality of axial injectors closer to the axis of rotation.
 20. The system of claim 17, wherein the plurality of axial injectors comprises an equal radial height, and a first injector of the plurality of axial injectors further away from the axis of rotation comprises a greater effective area for injection of the airflow into the airflow passage than a second injector of the plurality of axial injectors closer to the axis of rotation. 